Modern gas turbine engines are comprised of three major sections or components which function together to produce thrust for aircraft propulsion. In the compressor section, incoming ambient air is compressed and thus heated by a number of stages of rotating blades and stationary vanes. In the initial stages of the compressor the blades are generally made of titanium alloys, and in the later stages where temperatures are higher, the blades are generally made of iron or nickel base alloys. The compressed air may be heated to 1200.degree. to 1400.degree. F. at the last stage of compression, where it is passed on to the combustor where fuel is injected and burned. The hot gases exiting the combustor may be about 2400.degree. F., and are directed upon the first stage vane and blade of the turbine section. In the turbine section, comprised of a number of stages of rotating blades and stationary vanes, the actual work is extracted from the hot, compressed gases that turn the turbine which is connected to drive the earlier compressor section. A significant portion of the engine thrust comes from the large fan section at the front of the engine, which takes in ambient air and thrusts it backwards at a high velocity. The fan is also driven by the turbine section.
In the compressor, the early stages or the low compressor section are comprised of titanium alloy blades that rotate at high speed. The blades are designed such that their tips are very close to a stationary seal ring. The purpose of the close gap is to minimize gas leakage and to allow the pressure of the air to increase from one stage to the next. Narrow tip to seal gaps lead to higher engine efficiency and greater power output. If the gap is too narrow, there is the possibility of a rub between the tip and the seal. This can occur, for example, when the engine is started or if the pilot advances the throttle for more power. In these cases the blade can heat up faster than the surrounding case and through thermal expansion become longer and thus rub the seal ring. There are likely other mechanisms that also cause rubs. When the titanium alloy blade rubs the seal, the friction can be very high and the blade tip can heat up quickly to temperatures where the hot titanium can actually burn or oxidize with a further great liberation of heat. These situations are essentially titanium fires, and if left unchecked could damage the engine. Accordingly, a coating on the tip of these titanium blades is applied to separate the bare titanium from the seal material if a rub should occur.
In the turbine, the early stages of the high pressure turbine section are generally comprised of nickel base superalloy blades that rotate at high speed. These blades are also designed such that their tips are very close to a stationary seal ring. The purpose of the close gap is to minimize gas leakage and to allow the pressure of the air to do work against the turbine blades, causing them to rotate. Narrow tip to seal gaps lead to higher engine efficiency and greater power output. If the gap is too narrow, there is the possibility of a rub between the tip and the seal. As stated above, this can occur, for example, when the engine is started or if the pilot advances the throttle for more power. In these cases the blade can heat up faster than the surrounding case and through thermal expansion become longer and thus rub the seal ring. There are likely other mechanisms that also cause rubs. Typically, when a bare superalloy blade tip rubs against a bare cast superalloy seal, then the blade tip is worn back. In an improved design, the seal is coated with a material that is more rub tolerant than the cast seal material, and the seal takes a more significant fraction of the wear and the blade tip is less worn. However, that situation is still not ideal and coatings for blade tip are desired that reduce tip wear even more.
As engine temperatures are increased in a search for higher efficiency of operation, the metallic seal coatings suffer oxidation and some manufacturers are looking to ceramic seal coatings. In that case, the demands on a wear resistance blade tip coating increase even more. In a further progression of tip treatments, a composite layer of cubic boron nitride (CBN) embedded in a nickel or nickel alloy matrix is placed on the tip. This allows the tip coating to cut or grind into the seal material in a rub situation. However, this composite coating is difficult and expensive to apply to blade tips such as titanium blade tips.
U.S. Pat. No. 5,059,095 discloses a turbine rotor blade tip coated with alumina-zirconia ceramic. Specifically, the ceramic layer consisting of a combination of aluminum oxide and zirconium oxide or at least partially stabilized zirconium oxide. U.S. Pat. No. 5,073,433 discloses a thermal barrier coating comprising zirconia partially stabilized by yttria and having a substantially homogeneous dispersion of vertical macrocracks throughout the coating to improve its thermal fatigue.
It is an object of the present invention to provide a blade for a gas turbine engine whose tip segment is coated with a layer of a zirconium-based oxide having a plurality of vertical cracks and which has good rub tolerance when contacting a seal material such as a bare cast superalloy.
It is another object of the present invention to provide a stabilized zirconia coating for the tip portion of blades for a turbine engine having a plurality of macrocracks in the surface of the coating.
It is another object of the present invention to provide a turbine or compressor blade on which the tip of the blade has a zirconium based oxide coating with embedded particles more abrasive than zirconia and having a plurality of macrocracks throughout the coating.
It is another object of the present invention to provide a process for producing a coating for the tip of turbine and compressor blades having good rub tolerance characteristics.